Integrated Wiring System for Composite Structures

ABSTRACT

A method for manufacturing a composite part. Layers of composite material are cured to form the composite part. A primer is depicted on a surface of the composite part. A group of conductive elements is deposited on the primer to form an electronic device on the primer.

BACKGROUND INFORMATION

1. Field

The present disclosure relates generally to aircraft and, in particular,to electrical wiring in aircraft. Still more particularly, the presentdisclosure relates to a method and apparatus for depositing conductivematerials on composite structures used for aircraft.

2. Background

Manufacturing aircraft is a complex process. Thousands of components maybe manufactured and assembled to form an aircraft. These componentsinclude structural components, aircraft systems, control devices,passenger seating, storage compartments, and other suitable components.Some of these components include one or more composite materials.

To supply electricity to various aircraft systems, electrical wiring isinstalled throughout the aircraft. This electrical wiring may be runalong the entire fuselage of the aircraft and supply different types ofdevices within the aircraft. For instance, electrical wiring may supplya control surface, a display device, a light, an appliance, and othertypes of devices. Electrical wiring also may be run through each wing tosupply devices on the wing with power.

During installation, electrical wires are often mounted to aircraftstructures in bundles. These bundles comprise a plurality of individualwires which transmit electrical power.

A number of additional components are needed to secure these wirebundles to the aircraft. For example, without limitation, connectors,disconnect panels, ties, fasteners, clamps, standoffs, spanner bars,brackets, spacers, and other components may be needed to hold the wirebundles in place. Each of these components may be manufacturedseparately and then installed in the aircraft by an operator.

To properly install the wire bundles, an operator may drill holes in anaircraft structure before inserting components used to hold the wirebundles. For example, an operator may drill holes in a composite skinpanel to insert a fastener and a bracket to hold a wire bundle. Theseholes may be drilled at certain intervals for the brackets such thatsagging of the wiring is reduced. An operator also may have to assemblethe wire bundles using ties to hold the individual electrical wirestogether.

Performing these manufacturing and installation operations, as wellothers, may take more time than desired. As a result, the cost ofmanufacturing an aircraft may be increased. Additionally, manufacturingthe wiring bundles, as well as the components used to secure thesebundles to the aircraft, may add more weight and complexity to theaircraft than desired. Therefore, it would be desirable to have a methodand apparatus that take into account at least some of the issuesdiscussed above, as well as other possible issues.

SUMMARY

In one illustrative embodiment, a method for manufacturing a compositepart is presented. Layers of composite material are cured to form thecomposite part. A primer is deposited on the composite part. A group ofconductive elements is deposited on the primer to form an electronicdevice.

In another illustrative embodiment, an apparatus comprises a compositepart for an aircraft, a primer deposited on a surface of the compositepart, and an electronic device. The electronic device comprises a groupof conductive elements deposited on the primer. Power is supplied to adevice connected to the composite part through current flowing throughthe group of conductive elements.

In yet another illustrative embodiment, another method for manufacturinga composite part is presented. Layers of composite material are laid upfor the composite part. The layers of composite material are cured toform the composite part. A group of conductive elements is deposited ona layer of material using a direct-write process such that an electronicdevice is formed on the layer of material. The layer of material havingthe group of conductive elements is co-bonded with the composite part.Power is supplied to a device connected to the composite part throughcurrent flowing through the group of conductive elements.

The features and functions can be achieved independently in variousembodiments of the present disclosure or may be combined in yet otherembodiments in which further details can be seen with reference to thefollowing description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the illustrativeembodiments are set forth in the appended claims. The illustrativeembodiments, however, as well as a preferred mode of use, furtherobjectives and features thereof, will best be understood by reference tothe following detailed description of an illustrative embodiment of thepresent disclosure when read in conjunction with the accompanyingdrawings, wherein:

FIG. 1 is an illustration of an aircraft in accordance with anillustrative embodiment;

FIG. 2 is an illustration of a manufacturing environment in the form ofa block diagram in accordance with an illustrative embodiment;

FIG. 3 is an illustration of a block diagram of a wiring system inaccordance with an illustrative embodiment;

FIG. 4 is an illustration of a skin panel for a wing in accordance withan illustrative embodiment;

FIG. 5 is an illustration of a direct-write tool depositing conductivematerial in accordance with an illustrative embodiment;

FIG. 6 is an illustration of a skin panel with an electronic device inaccordance with an illustrative embodiment;

FIG. 7 is an illustration of a skin panel of a wing with an electronicdevice in accordance with an illustrative embodiment;

FIG. 8 is an illustration of a finished skin panel with a wiring systemin accordance with an illustrative embodiment;

FIG. 9 is an illustration of a wing for an aircraft in accordance withan illustrative embodiment;

FIG. 10 is an illustration of a layer of material for co-bonding with askin panel in accordance with an illustrative embodiment;

FIG. 11 is an illustration of a layer of material with an electronicdevice in accordance with an illustrative embodiment;

FIG. 12 is an illustration of a layer of material co-bonded with a skinpanel in accordance with an illustrative embodiment;

FIG. 13 is an illustration of a skin panel with a layer of materialhaving an electronic device in accordance with an illustrativeembodiment;

FIG. 14 is an illustration of a finished skin panel for a verticalstabilizer in accordance with an illustrative embodiment;

FIG. 15 is an illustration of a vertical stabilizer for an aircraft inaccordance with an illustrative embodiment;

FIG. 16 is an illustration of an interior panel with an electronicdevice in accordance with an illustrative embodiment;

FIG. 17 is an illustration of an interior panel installed in an aircraftin accordance with an illustrative embodiment;

FIG. 18 is an illustration of a flowchart of a process for manufacturinga composite part in accordance with an illustrative embodiment;

FIG. 19 is an illustration of a flowchart of a process for forming anelectronic device on a composite part in accordance with an illustrativeembodiment;

FIG. 20 is an illustration of a flowchart of a process for manufacturinga composite part in accordance with an illustrative embodiment;

FIG. 21 is an illustration of an aircraft manufacturing and servicemethod in the form of a block diagram in accordance with an illustrativeembodiment; and

FIG. 22 is an illustration of an aircraft in the form of a block diagramin which an illustrative embodiment may be implemented.

DETAILED DESCRIPTION

The illustrative embodiments recognize and take into account one or moredifferent considerations. For example, the illustrative embodimentsrecognize and take into account that it may be desirable to reduce thetime needed to install electrical wiring in an aircraft. For example,the illustrative embodiments recognize and take into account thatcountless labor hours are spent installing wire bundles throughout theaircraft. The illustrative embodiments recognize and take into accountthat an automated wiring process may reduce the time spent installingthese wire bundles.

The illustrative embodiments also recognize and take into account thatit may be desirable to reduce the weight of an aircraft to increase itsaerodynamic performance, fuel efficiency, and other performanceparameters. For instance, the illustrative embodiments recognize andtake into account that securing wiring bundles in an aircraft requiresuse of various accessory structures such as clamps, ties, and brackets.These accessory structures may add undesired weight to the aircraft.

The illustrative embodiments further recognize and take into accountthat it may be desirable to ensure structural integrity of compositeaircraft structures that are modified during installation of wirebundles in the aircraft. As an example, inspection is performed toensure that composite structures that were drilled, milled, or otherwisealtered during installation of the wire bundles still have a desiredlevel of structural integrity. This inspection may lead to reworking ordiscarding of composite aircraft structures, which increases the timeneeded to manufacture the aircraft.

Thus, the illustrative embodiments provide a method and apparatus formanufacturing a composite part with conductive elements. The compositepart may be configured for use in an aircraft. In an illustrativeembodiment, layers of composite material are cured to form the compositepart. A primer is deposited on the surface of the composite part aftercuring. A group of conductive elements is deposited on the primer suchthat an electronic device is formed on the primer on the surface of thecomposite part.

Referring now to the figures and, in particular, with reference to FIG.1, an illustration of an aircraft is depicted in accordance with anillustrative embodiment. In this depicted example, a perspective view ofaircraft 100 is shown.

As depicted, aircraft 100 has wing 102 and wing 104 attached to body106. Aircraft 100 includes engine 108 attached to wing 102 and engine110 attached to wing 104.

In this illustrative example, body 106 has tail section 112. Horizontalstabilizer 114, horizontal stabilizer 116, and vertical stabilizer 118are attached to tail section 112 of body 106.

Aircraft 100 has composite skin 120 in this illustrative example.Composite skin 120 is formed from layers of composite material. Theselayers of composite material have been laid up and cured upon exposureto temperature and pressure to form panels for composite skin 120. Inaddition to composite skin 120, other types of composite structures maybe arranged within the interior of aircraft 100. Skin panel 122 on wing104 and skin panel 124 on vertical stabilizer 118 are examples of panelsthat form composite skin 120.

An illustrative embodiment may be implemented in aircraft 100 to provideelectrical power to various devices used for aircraft 100. Specifically,a wiring system may be deposited on portions of composite skin 120 toprovide electrical power to these devices. A wiring system also may bedeposited on other composite structures within aircraft 100.

Turning next to FIG. 2, an illustration of a manufacturing environmentis depicted in the form of a block diagram in accordance with anillustrative embodiment. In this depicted example, manufacturingenvironment 200 is an environment in which wiring system 202 is formedfor mobile platform 204. Wiring system 202 in mobile platform 204 may beused to supply power 206 to device 208.

In this illustrative example, mobile platform 204 takes the form ofaircraft 210. Aircraft 100 in FIG. 1 is an example of a physicalimplementation for aircraft 210 shown in block form in this figure.

As depicted, wiring system 202 includes a number of differentcomponents. As used herein, a “number of” items may be one or moreitems. In this illustrative example, a number of components is one ormore components. Wiring system 202 includes composite part 212 andelectronic device 214 in this illustrative example.

As illustrated, composite part 212 is a structure formed from layers ofcomposite material 216. Each of the layers of composite material 216 mayinclude reinforcing fibers bound in polymer resin matrix. In thisillustrative example, the layers of composite material 216 include areinforcing material (e.g., fibers in sheets). These sheets may be inthe form of tapes, fabrics or take other suitable forms. In thisdepicted example, the fibers and resins are arranged and cured to formcomposite part 212.

Layers of composite material 216 may be formed of layers of compositeprepreg 218, in which the reinforcing material (e.g., fibers) has beeninfused or preimpregnated with resin. Layers of composite prepreg 218may be laid up on a tool and cured to form composite part 212. In otherillustrative examples, the resin may be infused into the reinforcingmaterial after layers of composite material 216 have been laid up toform the shape for composite part 212.

In this depicted example, composite part 212 is a cured composite part.In other words, layers of composite material 216 have already beenhardened. No additional curing process is needed for composite part 212to be used in a platform such as aircraft 100 in FIG. 1.

In this illustrative example, composite part 212 is shaped to take theform of aircraft part 220. Aircraft part 220 is a structure configuredfor use in aircraft 210. In some examples, composite part 212 may beshaped to take the form of an aircraft part 220 such as a skin panel, aninterior panel, a stringer, a frame, a spar, a wing, a winglet, afuselage, an empennage, a control surface, and other aircraft parts towhich or across which it may be desirable for wiring to be routed. Insome examples, composite part 212 is shaped to take the form of anaircraft part as described above before the electronic device 214 isprovided on composite part 212. In some examples, composite part 212 mayhave a complex shape, which may include one or more contoured surfaces.

As illustrated, composite part 212 has surface 222. Surface 222 may bean exterior surface of composite part 212. For instance, surface 222 maybe located on the most superficial layer of composite material used toform composite part 212. Surface 222 may be a contoured surface in someexamples.

In this illustrative example, electronic device 214 is deposited onsurface 222 of composite part 212, which may provide the composite part212 with additional functionality. For example, with electronic device214, composite part 212 may have functionality in addition to thecomposite part 212 functioning as a secondary or primary structure ofmobile platform 204.

In this particular example, electronic device 214 is a device configuredto send electrical signals to device 208. Specifically, electronicdevice 214 supplies power 206 to device 208. For example, electronicdevice 214 may be operable to supply power 206 to device 208 from powersource 207.

In this illustrative example, power source 207 may be a device locatedon composite part 212. As an example, power source 207 may be a solarcell located on a skin panel. In another example, power source 207 maybe located remote to composite part 212. In this case, power source 207may be electrically connected to composite part 212 in some manner. Forexample, power source 207 may be a battery system located remote toelectronic device 214 and electrically connected to electronic device214 via electronic traces at one or more points along composite part212. Power source 207 also may take other forms, depending on theparticular implementation.

As depicted, electronic device 214 comprises group of conductiveelements 224. As used herein, a “group of” items is more than one item.In this illustrative example, group of conductive elements 224 includesmore than one conductive element.

In this depicted example, a conductive element in group of conductiveelements 224 is a structure configured to conduct electrical signals.Group of conductive elements 224 comprise at least one of an electricaltrace, an interconnect, a wire, a transistor, an integrated circuit, aconductive connector, or other suitable types of conductive elements.

As used herein, the phrase “at least one of,” when used with a list ofitems, means different combinations of one or more of the listed itemsmay be used and only one of the items in the list may be needed. Theitem may be a particular object, thing, or category. In other words, “atleast one of” means any combination of items or number of items may beused from the list, but not all of the items in the list may berequired.

For example, “at least one of item A, item B, and item C” may mean itemA; item A and item B; item B; item A, item B, and item C; or item B anditem C. In some cases, “at least one of item A, item B, and item C” maymean, for example, without limitation, two of item A, one of item B, andten of item C; four of item B and seven of item C; or some othercombination.

As illustrated, group of conductive elements 224 comprises conductivematerial 226. Conductive material 226 is a type of material that allowsa flow of electrical signals in one or more directions. Conductivematerial 226 may include a metal, a metal alloy, or some other type ofconductor. Specifically, conductive material 226 may be selected from atleast one of copper, copper alloy, carbon, graphene, titanium, nickel,silver, or other suitable conductors.

In this illustrative example, power 206 is supplied to device 208connected to composite part 212 through current 228 flowing throughgroup of conductive elements 224. In one illustrative example,conductive traces are deposited on primer 221 extending from one end ofcomposite part 212 to another end of composite part 212 to form a powerrail for supplying power 206 across composite part 212.

In this depicted example, device 208 is an object configured to usepower 206 to perform operations in aircraft 210. Device 208 may be, forexample, without limitation, a light, an appliance, a control system, asensor, a display device, a computer, a field replaceable unit (FRU), aninflight entertainment system (IFE), a graphical indicator, a beacon, anemergency device, a door system, or some other suitable mechanism.

As an example, when device 208 is a light on a wing, electronic device214 may be deposited on surface 222 of the wing to supply power 206 tothe light. As another example, when device 208 includes interiorlighting in the passenger cabin of aircraft 210, electronic device 214may be deposited on surface 222 of an interior panel of the passengercabin to supply power 206 to the interior lighting. In still anotherillustrative example, when device 208 is a sensor, electronic device 214may be located on surface 222 of a skin panel to supply power 206 to thesensor to detect icing conditions for aircraft 210.

In this illustrative example, electronic device 214 may be depositedonto composite part 212 using deposition system 230. In particular,group of conductive elements 224 in electronic device 214 are depositedonto primer 221 applied to surface 222 of composite part 212 usingdeposition system 230.

In this illustrative example, an item is “deposited” onto a substrate,such as surface 222 of composite part 212 with primer 221, when it isadded to the substrate. In this manner, deposition of group ofconductive elements 224 is the “addition” or “application” of conductiveelements to the surface of composite part 212. In some cases, resistiveelements or other structures may be deposited on surface 222 in additionto group of conductive elements 224.

As depicted, deposition system 230 includes various componentsconfigured to apply group of conductive elements 224 to composite part212. For instance, deposition system 230 may include at least one of arobotic device, a sprayer, a brush, a gantry, a nozzle, a coating tool,a gun, a plasma spraying system, and other suitable types of tools.Deposition system 230 also may include a power source, a control device,a heating system, and other accessory components.

In this depicted example, deposition system 230 deposits group ofconductive elements 224 using direct-write process 232. In thisillustrative example, direct-write process 232 may comprise the additiveprocess of depositing only required material directly to a substrate.

For instance, direct-write process 232 comprises depositing onlyconductive material 226 on composite part 212. Direct-write process 232may differ from standard electronics processing techniques, which mayrequire that the substrate be completely covered with the desiredmaterial and that excess material be removed. This removal may bereferred to as “etching” in some illustrative examples.

In this illustrative example, direct-write process 232 may be selectedfrom one of a thermal plasma spray, a nano-particle ink-jet process,screen-printing, an atomized jetted ink process, kinetic metallization,and other suitable types of direct-write processes. With direct-writeprocess 232, no etching is needed to remove excess material from surface222 of composite part 212. For example, when using a thermal plasmaspray, conductive material 226 is directly sprayed in a desiredconfiguration on top of primer 221 and no etching is performed.

In this depicted example, primer 221 is deposited onto surface 222 toprepare surface 222 to receive conductive material 226. For instance,primer 221 may be deposited on surface 222 prior to the application ofconductive material 226 to promote adhesion of conductive material 226.In an illustrative example, primer 221 promotes adhesion bysubstantially reducing the spread of conductive material 226 toundesired areas on surface 222. In other words, primer 221 preventsexcess conductive material 226 from running.

Primer 221 also may provide other functionalities. For instance, primer221 may provide corrosion resistance, enhanced durability for conductivematerial 226, and increased structural load handling capabilities, amongothers. In further examples, primer 221 may provide an insulating layerbetween conductive material 226 and the composite part 212.

In this illustrative example, primer 221 is applied directly to surface222 in a manner that results in a uniform surface for depositingconductive material 226. In an illustrative example, a “uniform” surfacerefers to a surface that has substantially the same characteristicsacross its entirety. This uniform surface may be the portion of surface222 to which conductive material 226 is being applied. In some examples,this uniform surface may be a substantially smooth surface forapplication of conductive material 226. In another example, this uniformsurface may be a sanded surface. With a uniform surface, conductivematerial 226 may be deposited on primer 221 evenly.

Primer 221 is applied to surface 222 of composite part 212 using directwrite process 232. As an example, primer 221 may be sprayed directlyonto surface 222 of composite part 212 using thermal plasma spray.

In an illustrative example, primer 221 is deposited on surface 222 ofcomposite part 212 in one or more layers. Each layer of primer 221 maycomprise one or more materials. These materials may be selected from atleast one of metallic material 223, ceramic material 225, or some othersuitable type of material.

In this depicted example, when applying primer 221, a layer of metallicmaterial 223 is sprayed on surface 222 of composite part 212. A layer ofceramic material 225 is then sprayed on top of the layer of metallicmaterial 223. Additional layers of material may be deposited before orafter the layer of metallic material 223 and the layer of ceramicmaterial 225. In is illustrative example, conductive material 226 isthen deposited on layer of ceramic material 225 of primer 221.

Deposition system 230 may use a pre-planned electronic architecture todeposit conductive material 226 on primer 221 to form electronic device214 for composite part 212. For example, parameters 234 may be used todeposit conductive material 226 such that electronic device 214functions as desired.

As an example, an operator may input parameters 234 to controldeposition of conductive material 226. Examples of parameters 234include signal paths, orientations for the conductive elements,connection sites, dimensions, spacing, and other suitable parameters.Deposition system 230 may then use parameters 234 to deposit conductivematerial 226 in a desired manner.

In some illustrative examples, additional layers of material may beadded to surface 222 of composite part 212 prior to, during, or afterdeposition of primer 221 and conductive material 226. For example,primer 221 may comprise additional dielectric material 236. Further,protective material 241 may be added to composite part 212. Thesematerials may be sprayed onto composite part. 212, bonded to compositepart 212, or attached to composite part 212 in some other manner.

In this illustrative example, dielectric material 236 may be added tosurface 222 prior to formation of group of conductive elements 224.Specifically, dielectric material 236 may be part of primer 221 sprayedonto surface 222. For instance, a number of layers 240 of dielectricmaterial 236 may be sprayed on surface 222 under the layer of metallicmaterial 223. In other examples, a number of layers of dielectricmaterial 236 may be added after layer of metallic material 223. One ormore layers of dielectric material 236 may be added in addition to thelayer of ceramic material 225. In some cases, dielectric material 236may be added during or after depositing conductive material 226.Dielectric material 236 may be applied using direct-write process 232.

Dielectric material 236 may be a type of polymer configured to provideelectrical insulation for group of conductive elements 224 in thisillustrative example. For example, dielectric material 236 insulateseach of group of conductive elements 224 from one another. In anotherillustrative example, dielectric material 236 insulates group ofconductive elements 224 from surface 222 of composite part 212. In stillother illustrative examples, dielectric material 236 insulates group ofconductive elements 224 from metal fasteners and other structuresattached to composite part 212.

Dielectric material 236 may include a material selected from one offiberglass, direct-write deposited ceramic or polymer, a high solidspolymeric primer or paint, polymeric films such as polyimide,polyethylene terephthalate (PET), polyether ether ketone (PEEK),polyvinyl fluoride, and other suitable types of dielectric material.

In this depicted example, protective material 241 is a type of materialconfigured to shield group of conductive elements 224 from theenvironment around composite part 212. For example, protective material241 may protect group of conductive elements 224 from corrosion, weatherconditions, electromagnetic effects, other environmental conditions, ora combination thereof.

In this illustrative example, protective material 241 provideselectrical isolation from the environment. Protective material 241 alsoextends the life of group of conductive elements 224. Protectivematerial 241 also increases the safety of composite part 212 withconductive elements 224.

As depicted, a number of layers 242 of protective material 241 may beplaced on top of group of conductive elements 224 after direct-writeprocess 232 is completed. Protective material 241 may include a materialselected from one of a polymer, a polytetrafluoroethylene (Teflon) tape,ceramic, paint, a primer, and other suitable types of materials andcombinations of materials. The thickness of layers 242 of protectivematerial 241 may depend on the properties of the material.

Wiring system 202 provides power 206 to device 208 in aircraft 210without the need to install wiring in aircraft 210. Instead, electronicdevice 214 with group of conductive elements 224 may be depositeddirectly onto surface 222 of a cured composite part using direct-writeprocess 232.

With reference now to FIG. 3, an illustration of a block diagram of awiring system is depicted in accordance with an illustrative embodiment.In this illustrative example, an alternative embodiment of wiring system202 in manufacturing environment 200 is shown.

As depicted, layer of material 300 is co-bonded with surface 222 ofcomposite part 212. Specifically, layer of material 300 and surface 222of composite part 212 are co-bonded to form joint 302. Layer of material300 is co-bonded with surface 222 of composite part 212 after compositepart 212 has been trimmed, formed into a desired shape, and its surfaceprepared for co-bonding.

In this illustrative example, “co-bonding” is the process of joining twocomponents together in which one of the components is uncured and theother component is pre-cured. In other words, unlike co-curing, wheretwo components are cured at substantially the same time, co-bondinginvolves simultaneously curing one part and bonding it to another fullycured part. In this depicted example, composite part 212 is fully curedbefore joint 302 is formed between layer of material 300 and compositepart 212.

Layer of material 300 comprises thermoplastic film 306 in thisillustrative example. Thermoplastic film 306 is a polymer selected fromone of polyvinyl fluoride, polyamide, polyether ether ketone (PEEK),polyetherketoneketone (PEKK), and other suitable types of materials.

In some cases, adhesive 304 may be used to join layer of material 300and composite part 212. Adhesive 304 may be an epoxy or some othermaterial configured to bond layer of material 300 with surface 222 ofcomposite part 212.

In this depicted example, electronic device 214 is deposited on layer ofmaterial 300. Instead of depositing conductive material 226 directly onsurface 222 of composite part 212, deposition system 230 depositsconductive material 226 on layer of material 300 to form group ofconductive elements 224. Layer of material 300 is then co-bonded withcomposite part 212 such that electronic device 214 can supply power 206to device 208 as shown in block form in FIG. 2.

The illustration of wiring system 202 and the components within wiringsystem 202 in FIG. 2 and FIG. 3 are not meant to imply physical orarchitectural limitations to the manner in which an illustrativeembodiment may be implemented. Other components in addition to or inplace of the ones illustrated may be used. Some components may beoptional. Also, the blocks are presented to illustrate some functionalcomponents. One or more of these blocks may be combined, divided, orcombined and divided into different blocks when implemented in anillustrative embodiment.

Mobile platform 204 in FIG. 2 may take other forms other than aircraft210. For example, mobile platform 204 may be a surface ship, a tank, apersonnel carrier, a train, a spacecraft, a space station, a satellite,a submarine, an automobile, and other types of mobile platforms.

In some cases, direct-write process 232 may be implemented to formlayers of conductive elements on top of surface 222 of composite part212. For example, primer 221 may be applied to surface 222 and a firstelectronic device may be formed on primer 221. Additional primer 221 maybe applied to the first electronic device sufficient to insulate thefirst electronic device. A second electronic device may be formed on topof this primer. Additional primer may be added on top of the secondelectronic device, and so forth, to form layers of electronic devicesfor composite part 212.

As another example, additional layers of material may be added betweensurface 222 of composite part 212 and layer of material 300 shown inFIG. 3. Protective layers may be added after electronic device 214 isdeposited onto layer of material 300. These protective layers may beadded prior to co-bonding with composite part 212 or after co-bondinghas been completed.

In yet another illustrative example, one or more additional electronicdevices may be embedded within layers of composite material 216. One ormore additional electronic devices also may be embedded in layer ofmaterial 300.

In still another illustrative example, an electronic device may bedeposited on surface 222 of composite part 212 and another electronicdevice may be deposited on layer of material 300. In this case, thesedevices may send electrical signals to one another, supply power 206 todifferent devices, or perform in some other manner.

As another illustrative example, electronic device 214 may comprise abus. When electronic device 214 comprises a bus, electronic device 214sends a plurality of signals back and forth between devices. The bus maybe a power bus, a data bus, or perform some other function.

In yet another illustrative example, layer of material 300 and compositepart 212 may be joined after both components have been previously cured.For example, deposition system 230 first uses direct-write process 232to form group of conductive elements 224 on layer of material 300. Layerof material 300 is then cured. Finally, layer of material 300 with groupof conductive elements 224 is bonded to composite part 212. The twocomponents may be bonded using an adhesive, such as adhesive 304.

FIGS. 4-7 illustrate a process for depositing a group of conductiveelements onto a surface of a skin panel. An enlarged view of skin panel122 for wing 104 of aircraft 100 from FIG. 1 is shown in FIGS. 4-7. InFIG. 8, skin panel 122 is shown installed in aircraft 100. Skin panel122 is an example of a physical implementation for composite part 212shown in block form in FIG. 2.

In FIGS. 4-7, conductive elements are deposited on the surface of skinpanel 122 that faces the exterior of aircraft 100 in FIG. 1. However,the process shown in FIGS. 4-8 is applicable to all surfaces of skinpanel 122, including surfaces that face the interior of aircraft 100.For example, conductive elements 602 shown in FIGS. 6-7 may run throughthe interior of wing 104 on skin panel 122.

Turning to FIG. 4, an illustration of a skin panel for a wing isdepicted in accordance with an illustrative embodiment. In this depictedexample, skin panel 122 is shown in manufacturing environment 400 priorto being installed in wing 104. Manufacturing environment 400 is anexample of a physical implementation for manufacturing environment 200shown in block form in FIG. 2. A number of processes may be performed onskin panel 122 prior to installation in wing 104.

In this illustrative example, skin panel 122 comprises compositematerial 401. Composite material 401 may include layers of compositeprepreg that have been cured. In this manner, skin panel 122 is a curedpart that has already been trimmed to have a desired shape.

As depicted, deposition system 402 is positioned relative to skin panel122. Deposition system 402 includes robotic arm 404 and end effector406.

End effector 406 is configured to deposit primer 412 on surface 408 ofskin panel 122. In this illustrative example, primer 412 includes alayer of metallic material and a layer of ceramic material on the layerof metallic material.

Additional layers of dielectric material (not shown in this view) alsomay be added under or on top of these two layers. Primer 412 providesinsulation, durability, and other characteristics in this illustrativeexample.

In FIG. 5, an illustration of a direct-write tool depositing conductivematerial is depicted in accordance with an illustrative embodiment. Endeffector 406 takes the form of a direct-write tool in this illustrativeexample.

As shown, end effector 406 is configured to deposit conductive material500 on primer 412. Primer 412 is transparent in this view such that thedeposition of conductive material 500 may be seen more clearly.Conductive material 500 may be desired on leading edge 502 of skin panel122 such that power may be supplied to a light at the distal end of wing104.

Turning to FIG. 6, an illustration of a skin panel with an electronicdevice is depicted in accordance with an illustrative embodiment. Inthis depicted example, electronic device 600 has been formed on surface408 of skin panel 122 from FIG. 5.

As illustrated, electronic device 600 comprises conductive elements 602.Deposition system 402 has deposited conductive material 500 on primer412 to form conductive elements 602.

As depicted, deposition system 402 uses a direct-write technique todeposit conductive material 500 directly onto primer 412 of skin panel122 without the need for etching afterward. Skin panel 122 withelectronic device 600 may be part of a wiring system for aircraft 100 inFIG. 1.

Next, in FIG. 7, an illustration of a skin panel of a wing with anelectronic device is depicted in accordance with an illustrativeembodiment. In this example, end effector 406 shown in FIG. 6 has beenconfigured to spray protective material 700 on top of group ofconductive elements 602 in electronic device 600. Protective material700 shields group of conductive elements 602 from the environmentsurrounding skin panel 122.

In other illustrative examples, protective material 700 may be appliedin some other manner. For instance, a layer of protective material 700may be bonded to skin panel 122 on top of conductive elements 602.

FIG. 8 shows an illustration of a finished skin panel with a wiringsystem in accordance with an illustrative embodiment. In this depictedexample, skin panel 122 is ready to be installed in wing 104 shown inFIG. 1.

In FIG. 9, an illustration of a wing for an aircraft is depicted inaccordance with an illustrative embodiment. An enlarged view of wing 104for aircraft 100 from FIG. 1 is shown in this figure.

Skin panel 122 has been installed in wing 104. Light 900 is connected toskin panel 122 on wing 104 in this illustrative example. As depicted inthis example, electronic device 600 (not shown in this view) suppliespower to light 900 using conductive elements 602.

Although FIGS. 4-9 show the deposition of electronic device 600 on askin panel for a wing, electronic device 600 may be deposited in asimilar manner for other structures in aircraft 100. These structuresinclude a wing tip, a stringer, an interior panel, or other aircraftcomposite structures.

Electronic device 600 also may be used in conjunction with wiredsystems. For example, a portion of aircraft 100 may be wired usingphysical wire bundles and a portion of aircraft 100 may includeelectronic devices deposited in the manner described in FIGS. 4-9.

Connections for electronic device 600 may be made without substantialrework or modification of skin panel 122. For instance, a portion ofprotective material 700 may be removed to expose one or more ofconductive elements 602. Conductive elements 602 may then be connectedto a power source or other structure, without drilling holes in thecomposite part.

FIGS. 10-14 illustrate a process for co-bonding a layer of material witha skin panel. The layer of material comprises conductive elements. Anenlarged view of the layer of material is shown in FIGS. 10-14. Anenlarged view of skin panel 124 for vertical stabilizer 118 of aircraft100 from FIG. 1 is shown in FIGS. 12-14. In FIG. 15, skin panel 124 isshown installed in aircraft 100.

Layer of material 1000 is an example of a physical implementation oflayer of material 300 shown in block form in FIG. 3. Skin panel 124 isan example of a physical implementation for composite part 212 shown inblock form in FIG. 3.

Turning to FIG. 10, an illustration of a layer of material forco-bonding with a skin panel is depicted in accordance with anillustrative embodiment. In this depicted example, layer of material1000 is shown in manufacturing environment 400 prior to being co-bondedwith skin panel 124 for vertical stabilizer 118 in FIG. 1.

A number of processes may be performed on layer of material 1000 priorto co-bonding with skin panel 124. In this illustrative example, layerof material 1000 comprises thermoplastic film 1002. Thermoplastic film1002 takes the form of polyvinyl fluoride in this illustrative example.

In this depicted example, deposition system 402 is positioned relativeto layer of material 1000. End effector 406 deposits conductive material500 onto surface 1004 of layer of material 1000.

It may be desirable to co-bond layer of material 1000 having conductivematerial 500 to skin panel 124 such that power may be supplied to alight at the distal end of vertical stabilizer 118. A number of layersof dielectric material may have been placed on layer of material 1000 toprovide insulation in this illustrative example.

In FIG. 11, an illustration of a layer of material with an electronicdevice is depicted in accordance with an illustrative embodiment. Inthis depicted example, electronic device 1100 has been formed on surface1004 of layer of material 1000.

As illustrated, electronic device 1100 comprises conductive elements1102. Deposition system 402 has deposited conductive material 500 onsurface 1004 of layer of material 1000 to form conductive elements 1102.

Deposition system 402 uses a direct-write process in this illustrativeexample. In this manner, conductive material 500 is deposited withoutthe need of post-deposition etching. Electronic device 1100 withconductive elements 1102 may be part of a wiring system for aircraft 100in FIG. 1.

Next, in FIG. 12, an illustration of a layer of material co-bonded witha skin panel is depicted in accordance with an illustrative embodiment.In this depicted example, skin panel 124 has been previously cured andlayer of material 1000 is uncured.

As shown, layer of material 1000 is co-bonded to skin panel 124. Anoperator may use various co-bonding techniques to join the twostructures together. In this manner, electronic device 1100 is immovablyattached to surface 1200 of skin panel 124.

Turning now to FIG. 13, an illustration of a skin panel with a layer ofmaterial having an electronic device is depicted in accordance with anillustrative embodiment. In this example, end effector 406 spraysprotective material 1300 on top of group of conductive elements 1102 inelectronic device 1100. Protective material 1300 shields group ofconductive elements 1102 from the environment surrounding skin panel124.

FIG. 14 shows an illustration of a finished skin panel for a verticalstabilizer in accordance with an illustrative embodiment. In thisdepicted example, skin panel 124 is ready to be installed in verticalstabilizer 118 shown in FIG. 1.

In FIG. 15, an illustration of a vertical stabilizer for an aircraft isdepicted in accordance with an illustrative embodiment. In this depictedexample, an enlarged view of vertical stabilizer 118 for aircraft 100from FIG. 1 is shown.

Skin panel 124 has been installed in vertical stabilizer 118. Light 1500is connected to skin panel 124 on vertical stabilizer 118 in thisillustrative example. Electronic device 1100 (not shown in this view)supplies power to light 1500 using conductive elements 1102.

Referring next to FIG. 16, an illustration of an interior panel with anelectronic device is depicted in accordance with an illustrativeembodiment. In this depicted example, interior panel 1600 is shown afterconductive elements 1602 have been deposited.

Interior panel 1600 comprises composite material 1604. Interior panel1600 is another example of a physical implementation for composite part212 shown in block form in FIG. 2.

Interior panel 1600 is an example of a composite panel used in theinterior of aircraft 100 in FIG. 1. In particular, interior panel 1600may form part of the ceiling of a passenger cabin in aircraft 100.

In this depicted example, interior panel 1600 has side 1606 and side1608. Side 1608 is opposite side 1606. When installed in aircraft 100,side 1608 faces the interior of the passenger cabin.

As depicted, plurality of light sources 1610 and audio system 1612 arealso arranged on interior panel 1600. Plurality of light sources 1610may be light-emitting diodes (LEDs) or other illumination devices. Audiosystem 1612 may include a plurality of speakers in this illustrativeexample.

As shown in this view, conductive elements 1602 electrically couple atleast one of the plurality of light sources from the plurality of lightsources 1610, audio system 1612, or a combination thereof to a powersource. Conductive elements 1602 are used to power at least one ofplurality of light sources 1610 and audio system 1612 in thisillustrative example.

FIG. 17 shows an illustration of an interior panel installed in anaircraft in accordance with an illustrative embodiment. Interior panel1600 from FIG. 16 has been installed in cabin 1700. Cabin 1700 islocated in the interior of aircraft 100 shown in FIG. 1. Plurality oflight sources 1610 illuminate a portion of cabin 1700.

In this instance, plurality of light sources 1610 provide track lightingfor passengers in aircraft 100. Illumination of plurality of lightsources 1610 is provided using power supplied by conductive elements1602. Audio system 1612 provides sound for passengers in cabin 1700,using power supplied by conductive elements 1602.

The illustrations of electronic devices, conductive elements, andmanufacturing processes shown in FIGS. 4-17 are not meant to implyphysical or architectural limitations to the manner in which anillustrative embodiment may be implemented. Other components in additionto or in place of the ones illustrated may be used. Some components maybe optional.

The different components shown in FIGS. 4-17 may be illustrativeexamples of how components shown in block form in FIGS. 2-3 can beimplemented as physical structures. Additionally, some of the componentsin FIGS. 4-17 may be combined with components in FIGS. 2-3, used withcomponents in FIG. 2-3, or a combination of the two.

Further, although electronic device 600 and electronic device 1100 areshown and described as providing power to illumination devices, each ofthese devices may be configured to provide power to various other typesof devices. For instance, electronic device 600 may power a sensorlocated on wing 104.

As another example, electronic device 1100 may power a control surfacelocated on vertical stabilizer 118. Each of these electronic devicesalso may power multiple units in some illustrative examples.

With reference now to FIG. 18, an illustration of a flowchart of aprocess for manufacturing a composite part is depicted in accordancewith an illustrative embodiment. The process illustrated in FIG. 18 maybe implemented in manufacturing environment 200 in FIG. 2. The differentoperations may be implemented to form wiring system 202 in FIG. 2.

The process begins by curing layers of composite material to form acomposite part (operation 1800). Next, the process deposits a primer ona surface of the composite part (operation 1802). A group of conductiveelements is then deposited on the primer such that an electronic deviceis formed on the primer (operation 1803) with the process terminatingthereafter.

Referring next to FIG. 19, an illustration of a flowchart of a processfor forming an electronic device on a composite part is depicted inaccordance with an illustrative embodiment. The process illustrated inFIG. 19 may be implemented in manufacturing environment 200 bydeposition system 230 to form wiring system 202 in FIG. 2.

The process begins by depositing a layer of metallic material on thesurface of the composite part (operation 1900). As an example, a layerof metallic material may be sprayed on directly on the surface of thecomposite part using a thermal plasma spray.

Next, the process deposits a layer of ceramic material on the layer ofmetallic material (operation 1902). This layer of ceramic material alsomay be sprayed on using a thermal plasma spray. Optionally, a number ofadditional layers of material, such as dielectric material, may beapplied to at least one of the surface of the composite part, the layerof metallic material, or the layer of ceramic material.

Thereafter, the process deposits conductive material on the layer ofceramic material (operation 1904). As an example, the conductivematerial may be sprayed on the layer of ceramic material using a thermalplasma spray to form the group of conductive elements.

The process then applies a number of layers of protective material ontop of the conductive material (operation 1906), with the processterminating thereafter. The composite part is now ready for inspectionand installation in the aircraft.

In FIG. 20, an illustration of a flowchart of a process formanufacturing a composite part is depicted in accordance with anillustrative embodiment. The process illustrated in FIG. 20 may beimplemented in manufacturing environment 200 in FIG. 2. The differentoperations may be implemented to form wiring system 202.

The process begins by laying up layers of composite material for thecomposite part (operation 2000). Next, the process cures the layers ofcomposite material to form the composite part (operation 2002).

Thereafter, the process deposits a group of conductive elements on alayer of material using a direct-write process such that an electronicdevice is formed on the layer of material (operation 2004). The processthen positions the layer of material relative to a surface of thecomposite part (operation 2006).

Next, the process co-bonds a layer of material with the composite part(operation 2008) with the process terminating thereafter. The conductiveelements may now be used to supply power to various devices connected tothe composite part.

The flowcharts and block diagrams in the different depicted embodimentsillustrate the architecture, functionality, and operation of somepossible implementations of apparatuses and methods in an illustrativeembodiment. In this regard, each block in the flowcharts or blockdiagrams may represent at least one of module, a segment, a function, ora portion a combination thereof of an operation or step.

In some alternative implementations of an illustrative embodiment, thefunction or functions noted in the blocks may occur out of the ordernoted in the figures. For example, in some cases, two blocks shown insuccession may be executed substantially concurrently, or the blocks maysometimes be performed in the reverse order, depending upon thefunctionality involved. Also, other blocks may be added in addition tothe illustrated blocks in a flowchart or block diagram.

Illustrative embodiments of the disclosure may be described in thecontext of aircraft manufacturing and service method 2100 as shown inFIG. 21 and aircraft 2200 as shown in FIG. 22. Turning first to FIG. 21,an illustration of an aircraft manufacturing and service method isdepicted in the form of a block diagram in accordance with anillustrative embodiment. During pre-production, aircraft manufacturingand service method 2100 may include specification and design 2102 ofaircraft 2200 in FIG. 22 and material procurement 2104.

During production, component and subassembly manufacturing 2106 andsystem integration 2108 of aircraft 2200 in FIG. 22 takes place.Thereafter, aircraft 2200 in FIG. 22 may go through certification anddelivery 2110 in order to be placed in service 2112. While in service2112 by a customer, aircraft 2200 in FIG. 22 is scheduled for routinemaintenance and service 2114, which may include modification,reconfiguration, refurbishment, and other maintenance or service.

Each of the processes of aircraft manufacturing and service method 2100may be performed or carried out by a system integrator, a third party,an operator, or a combination thereof. In these examples, the operatormay be a customer. For the purposes of this description, a systemintegrator may include, without limitation, any number of aircraftmanufacturers and major-system subcontractors; a third party mayinclude, without limitation, any number of vendors, subcontractors, andsuppliers; and an operator may be an airline, a leasing company, amilitary entity, a service organization, and so on.

With reference now to FIG. 22, an illustration of an aircraft isdepicted in the form of a block diagram in which an illustrativeembodiment may be implemented. In this example, aircraft 2200 isproduced by aircraft manufacturing and service method 2100 in FIG. 21and may include airframe 2202 with plurality of systems 2204 andinterior 2206. Examples of systems 2204 include one or more ofpropulsion system 2208, electrical system 2210, hydraulic system 2212,and environmental system 2214. Any number of other systems may beincluded. Although an aerospace example is shown, different illustrativeembodiments may be applied to other industries, such as the automotiveindustry.

Apparatuses and methods embodied herein may be employed during at leastone of the stages of aircraft manufacturing and service method 2100 inFIG. 21. In particular, wiring system 202 from FIGS. 2-3 may be formedduring component and subassembly manufacturing 2106. For instance, groupof conductive elements 224 may be deposited on surface 222 of compositepart 212 used in aircraft 2200 during component and subassemblymanufacturing 2106.

As another example, layer of material 300 and composite part 212 may beco-bonded during component and subassembly manufacturing 2106. In yetanother illustrative example, wiring system 202 may be installed inaircraft 2200 during system integration 2108, routine maintenance andservice 2114, or some other stage of aircraft manufacturing and servicemethod 2100.

In one illustrative example, components or subassemblies produced incomponent and subassembly manufacturing 2106 in FIG. 21 may befabricated or manufactured in a manner similar to components orsubassemblies produced while aircraft 2200 is in service 2112 in FIG.21. As yet another example, one or more apparatus embodiments, methodembodiments, or a combination thereof may be utilized during productionstages, such as component and subassembly manufacturing 2106 and systemintegration 2108 in FIG. 21. One or more apparatus embodiments, methodembodiments, or a combination thereof may be utilized while aircraft2200 is in service 2112, during maintenance and service 2114 in FIG. 21,or a combination thereof. The use of a number of the differentillustrative embodiments may substantially expedite the assembly, reducethe cost of aircraft 2200, or both.

Thus, the illustrative embodiments provide a method and apparatus formanufacturing composite part 212 with group of conductive elements 224.Composite part 212 with conductive elements 224 form wiring system 202.In an illustrative embodiment, layers of composite material 216 arecured to form composite part 212. Primer 221 with layer of metallicmaterial 223 and layer of ceramic material 225 are deposited on surface222 of composite part 212. Group of conductive elements 224 is depositedon primer 221 such that electronic device 214 is formed on primer 221 onsurface 222 of composite part 212. In another illustrative embodiment,after curing and trimming, composite part 212, layer of material 300having group of conductive elements 224 is co-bonded with composite part212. Both embodiments provide a means to supply power 206 from powersource 207 to device 208 in this illustrative example.

With the use of an illustrative embodiment, a wiring system for anaircraft may be integrated into a composite part. The process fordepositing conductive elements onto the part may be automated using adirect-write process. The direct-write process allows deposition ofconductive elements to occur upstream in the manufacturing processwithout the need for etching or other time-consuming post-processingsteps. As a result, the time needed to manufacture wiring systems foraircraft may be reduced.

The illustrative embodiments also reduce the time needed to installwiring systems in aircraft. Instead of manufacturing each physicalwiring component, assembling those components into wire bundles, andinstalling the wire bundles in the aircraft, the use of an illustrativeembodiment allows operators to bypass those steps for some compositestructures.

In some cases, a wiring system in accordance with an illustrativeembodiment may be used throughout the aircraft. In other cases, such asystem may be used in conjunction with existing wiring systems usingwire bundles. The reduction in labor hours results in a reduction incost to manufacture the aircraft. Moreover, the illustrative embodimentsestablish a reliable and repeatable process for depositing conductiveelements on a composite structure.

In addition, when conductive elements are deposited on compositeaircraft parts, significant weight savings are realized. The need foraccessory structures, such as ties, clamps, and brackets, among others,is reduced. As a result, the use of an illustrative embodiment decreasesthe weight of the aircraft. This decrease in weight may increase theaerodynamic performance of the aircraft.

The illustrative embodiments also aid in ensuring the structuralintegrity of composite aircraft parts. Instead of drilling holes,countersinking those holes, and installing fasteners, electronic devicesare formed on the composite part without damaging the structure. In thismanner, inspection of each hole for cracks, delamination, and otherinconsistencies can be reduced or eliminated, resulting in additionaltime and cost savings.

The description of the different illustrative embodiments has beenpresented for purposes of illustration and description, and is notintended to be exhaustive or limited to the embodiments in the formdisclosed. Many modifications and variations will be apparent to thoseof ordinary skill in the art. Further, different illustrativeembodiments may provide different features as compared to otherdesirable embodiments. The embodiment or embodiments selected are chosenand described in order to best explain the principles of theembodiments, the practical application, and to enable others of ordinaryskill in the art to understand the disclosure for various embodimentswith various modifications as are suited to the particular usecontemplated.

What is claimed is:
 1. A method for manufacturing a composite partcomprising: curing layers of composite material to form the compositepart; depositing a primer on a surface of the composite part; anddepositing a group of conductive elements on the primer such that anelectronic device is formed on the composite part.
 2. The method ofclaim 1, wherein depositing the primer on the surface of the compositepart comprises: depositing a layer of metallic material on the surfaceof the composite part; and depositing a layer of ceramic material on thelayer of metallic material.
 3. The method of claim 1, wherein depositingthe primer on the surface of the composite part comprises: spraying alayer of metallic material on the surface of the composite part using athermal plasma spray; and spraying a layer of ceramic material on thelayer of metallic material using the thermal plasma spray.
 4. The methodof claim 3, wherein depositing the group of conductive elementscomprises: spraying conductive material on the layer of ceramic materialusing the thermal plasma spray to form the group of conductive elements.5. The method of claim 1, wherein depositing the group of conductiveelements on the primer comprises: depositing conductive traces on theprimer extending from one end of the composite part to another end ofthe composite part to form a power rail for supplying power across thecomposite part.
 6. The method of claim 1 further comprising: depositingthe group of conductive elements on the primer using a direct-writeprocess.
 7. The method of claim 6, wherein the direct-write process isselected from one of a thermal plasma spray, a nano-particle ink-jetprocess, screen-printing, an atomized jetted ink process, and kineticmetallization.
 8. The method of claim 1, wherein depositing the primeron the surface of the composite part comprises: applying a number oflayers of dielectric material to the surface of the composite part priorto depositing the group of conductive elements.
 9. The method of claim 1further comprising: applying a number of layers of protective materialon top of the group of conductive elements.
 10. The method of claim 1,wherein depositing the group of conductive elements comprises:depositing at least one of an electrical trace, an interconnect, a wire,a transistor, an integrated circuit, or a conductive connector.
 11. Themethod of claim 1, wherein the composite part is an aircraft part. 12.An apparatus comprising: a composite part for an aircraft; and a primerdeposited on a surface of the composite part; and an electronic devicecomprising a group of conductive elements deposited on the primer,wherein power is supplied to a device connected to the composite partthrough current flowing through the group of conductive elements. 13.The apparatus of claim 12, wherein the primer comprises: a layer ofmetallic material sprayed on the surface of the composite part using athermal plasma spray; and a layer of ceramic material sprayed on thelayer of metallic material using the thermal plasma spray.
 14. Theapparatus of claim 13, wherein conductive material is sprayed on theceramic material using the thermal plasma spray to form the group ofconductive elements.
 15. The apparatus of claim 14, wherein theconductive material is selected from at least one of copper, copperalloy, carbon, graphene, titanium, nickel, or silver.
 16. The apparatusof claim 12, wherein the composite part is selected from one of a skinpanel, an interior panel, a stringer, a frame, a spar, a wing, awinglet, a fuselage, an empennage, and a control surface.
 17. Theapparatus of claim 12, wherein the group of conductive elementscomprises at least one of an electrical trace, an interconnect, a wire,a transistor, an integrated circuit, or a conductive connector.
 18. Amethod for manufacturing a composite part comprising: laying up layersof composite material for the composite part; curing the layers ofcomposite material to form the composite part; depositing a group ofconductive elements on a layer of material using a direct-write processsuch that an electronic device is formed on the layer of material; andco-bonding the layer of material having the group of conductive elementswith the composite part, wherein power is supplied to a device connectedto the composite part through current flowing through the group ofconductive elements.
 19. The method of claim 18, wherein depositing thegroup of conductive elements comprises: spraying conductive material onthe layer of material using a thermal plasma spray to form the group ofconductive elements.
 20. The method of claim 18, wherein the layer ofmaterial comprises a thermoplastic film.